Hybrid composite structure having damped metallic fibers and method for making the same

ABSTRACT

A damped composite structure is formed from a matrix material and a plurality of shape memory alloy wire fibers held in the material matrix for damping the structure. The wire fibers may be embedded in a viscoelastic interlayer to increase damping of the structure. The wire fibers may be interspersed with reinforcement such as carbon fibers, in tows or in a mesh of fibers. The wire fibers have an inherent material loss factor greater than approximately 0.10, and may be formed from superelastic metal alloys, such as Ni—Ti, Cu—Zn—Al, Cu—Al—Ni, OR Cu—Al—Be.

TECHNICAL FIELD

This disclosure generally relates to composite material structures,especially those used in aerospace applications, and deals moreparticularly with a composite structure having embedded metallic fibersfor damping sound and vibration.

BACKGROUND

Noise and low frequency vibration may be created in commercial aircraftby a variety of sources. For example, exterior noise may be created byair passing over the aircraft's outer skin in a turbulent boundarylayer. Also, wing mounted engines may excite low frequency vibrationmodes of the aircraft's fuselage. This vibration energy may betransmitted through the fuselage by stringers and frames and enter intothe cabin interior as noise.

One solution for reducing noise and vibration involves the use ofdamping treatments such as “add-on” patches of viscoelastic materialswhich absorb and dampen noise and vibration. These add-on patches areessentially “peel and stick” components that are added to various partsof the aircraft, such as skins, frames and floor panels. A typicalcommercial aircraft may employ as many as 2500 to 3500 of these patches.While effective in reducing noise and vibration, the patches add weightto the aircraft and are labor intensive to install.

Accordingly, there is a need for a composite structure having inherentdamping qualities that reduce or eliminate the need for add-on dampingtreatments of the type mentioned above. Embodiments of the disclosureare intended to satisfy this need.

SUMMARY

The disclosed embodiments provide a composite structure possessinginherent sound damping that reduces sound and vibration before they aretransmitted to the interior of the cabin. The composite structure relieson the use of embedded metallic fibers formed of super-elastic metals(SEM) (also commonly referred to as shape memory alloys (SMAs)) to dampsound and vibration energy. The metallic fibers may be interspersedalong with high strength fibers such as carbon fibers in laminated pliesof synthetic resin. The metallic fibers undergo a reversible solid-statephase transformation from austenite to martensite when deformed orsubjected to reduced temperatures. Accordingly, at colder temperaturesnormally encountered by commercial aircraft at cruise altitudes, themetallic fibers are more deformable and therefore provide greaterinherent damping.

According to one disclosed embodiment, a damped composite structure isprovided, comprising a matrix material and a plurality of SEM wirefibers held in the matrix material for damping the structure. In oneform, the material matrix may optionally include a layer of resincontaining an interlayer of viscoelastic material in which a group ofwire fibers are embedded. In another form, the wire fibers may beinterspersed with reinforcing carbon fibers infused with a syntheticresin. The wire fibers may be interspersed in tows of reinforcingfibers, or arranged unidirectionaly in the form of tape or fabric alongwith the reinforcing fibers. The wire fibers may also be arranged as amesh held within a synthetic resin matrix forming a material ply. Thewire fibers may be formed from any of several super-elastic metalalloys, including Ni—Ti, Cu—Zn—Al, Cu—Al—Ni and Cu—Al—Be. The wirefibers may exhibit an inherent material loss factor greater thanapproximately 0.10.

According to another disclosed embodiment, a damped composite aircraftskin is provided, comprising: laminated plies of synthetic resin; afirst set of fibers in the resin for reinforcing the plies; and, asecond set of metallic fibers in the resin for damping noise andvibration, and which have an inherent material loss factor greater than0.10. The first and second sets of fibers may be arranged in tows, or inparallel rows within each ply of the resin. Some of the plies mayoptionally include a viscoelastic interlayer in which the metallicfibers in the second set are embedded.

In accordance with a method embodiment, a method is provided forfabricating a damped composite structure for aircraft, comprising thesteps of: introducing super-elastic metal alloy wire fibers into pliesof fiber reinforced synthetic resin; and, laminating the plies. Theplies may be laminated by laying up multiple fiber reinforced plies ofsynthetic epoxy resin having the wire fibers introduced therein,compacting the plies of the lay-up and then curing the compacted lay-up.The wire fibers may be introduced into the plies by forming tows ofreinforcing fibers containing the wire fibers. Alternatively, the wirefibers may be introduced to the plies by forming a prepreg tape thatincludes both wire fibers and reinforcing fibers.

Other features, benefits and advantages of the disclosed embodimentswill become apparent from the following description of embodiments, whenviewed in accordance with the attached drawings and appended claims.

BRIEF DESCRIPTION OF THE ILLUSTRATIONS

FIG. 1 is a perspective illustration of a commercial aircraft showingtypical sources of noise and vibration.

FIG. 2 is a cross sectional illustration of one embodiment of thecomposite structure having embedded damping.

FIG. 3 illustrates a temperature hysteresis curve of a typical shapememory alloy.

FIG. 4 is a cross sectional illustration of a typical tow of carbonfibers having interspersed wire fibers according to one embodiment.

FIG. 5 is a perspective illustration of a section of a ply employingmultiple layers of fiber reinforcement including interspersed wirefibers.

FIG. 6 is a cross sectional illustration of a layer of reinforcingfibers in which wire fibers are placed in the spaces between thereinforcing fibers.

FIG. 7 is a fragmentary, cross sectional illustration of anotherembodiment of the composite structure in which the wire fibers areembedded in plies alternating with plies containing reinforcing fibers.

FIG. 8 is a fragmentary, cross sectional illustration of one of theplies of the composite structure illustrated in FIG. 6, showing the wirefibers held in a resin matrix.

FIG. 9 is a plan illustration of an optional interlayer of viscoelasticmaterial containing an embedded mesh of wire fibers.

FIG. 10 is a sectional illustration showing the interlayer of FIG. 8held within a laminated resin ply.

FIG. 11 is a diagrammatic illustration of a line for fabricating prepregtape containing both reinforcing fibers and wire fibers.

FIG. 12 is a graph relating levels of damping with their effect onstiffness as a function of the volume percentage amount SEM wire fibersused in a composite structure.

DETAILED DESCRIPTION

Referring to FIG. 1, typical commercial and military aircraft 20 maygenerate noise outside the aircraft that is transmitted throughstringers and other structures (not shown) to an interior cabin withinthe aircraft's fuselage 22. This exterior sound and vibration,collectively referred to herein as “noise”, may originate from any ofseveral sources. For example, a primary source of exterior noise may bedue to air flowing over the aircraft 20 within the turbulent boundarylayer on the aircraft's skin. Also, in the case of wing-mounted engineaircraft, the engines may produce a buzz saw noise indicated at 26 thatis radiated directly to the fuselage 20, as well as shock cell noise 28from the engine exhaust as a result of a pressure differential betweenthe exhaust gasses and the ambient air. The shock cell noise 28 mayexcite low frequency vibration modes of the fuselage 22, resulting innoise within the cabin.

In accordance with embodiments of the disclosure, a composite structuregenerally indicated at 30 in FIG. 2 possesses inherent damping that mayreduce or eliminate noise within the cabin as a result of noise outsidethe fuselage 22. The composite structure 30 may be used in theconstruction of, without limitation, fuselage skins, wing skins, doorpanels and access panels, stiffening members, supports, struts, sparsand other structural members that may act to transmit noise to the cabininterior.

In accordance with the disclosed embodiments, the composite structure 30achieves noise damping by incorporating super-elastic wire fibers 46,sometimes referred to herein as metallic fibers formed of a shapeelastic metal alloy (SEM), also referred to shape memory alloys (SMA).SEM's are metal alloys that exhibit at least two unique properties:pseudo-elasticity and the shape memory effect. These two properties area result of a solid state phase change consisting of molecularrearrangement, which occurs in the SEM material. When mechanically orthermally stressed, the SEM material undergoes a reversible solid statephase transformation from austenite to martensite upon cooling (or bydeformation), and reverses the transformation upon heating or release ofthe oppose stress.

Martensite is a relatively soft and easily deformed phase of SEMmaterials which exists at lower temperature. Austenite is the strongerphase of SEM materials, which occurs at higher temperatures. SEMmaterials may be particularly useful for damping in aircraftapplications due to the fact that at colder temperature, it is moredeformable and softer in its martensite phase, and at highertemperatures it is stronger in its austenite phase. In its martensitephase, the inherent damping of the SEM material is an order of magnitudehigher than in its austenite phase.

The high damping of SEM materials results from its ability to transformmechanical energy (produced by an applied force, for example) intothermal energy in the form of heat dissipation. This energytransformation allows the SEM material to resist shock and also absorbvibration. The internal friction in the form of heat disperses theenergy between the different phases, or within the same phase(martensite, austenite).

Referring to FIG. 3, depending on the temperature and stress on thecomposite structure 30, the optimum SEM alloy for a particularapplication may be chosen with consideration given to the start andfinish of the martensite and austenite phases. In FIG. 3, thetemperatures at which these phases occur are designated as Ms(Martensite start), Mf (Martensite finish), As (Austenite start) and Af(Austenite finish). In other words, Ms denotes the temperature at whichthe structure begins to change from the austenite to the martensite uponcooling, while Mf is the temperature at which the transition isfinished. Similarly, As and Af are the temperatures at which the reversetransformation from martensite to austenite start and finish,respectively.

Generally, the better damping will occur in the low temperaturemartensite phase. However, desirable damping results may be obtained bydesigning a hybrid fiber composite structure 30 using SEM wire fibers 46that are in their transition phase. Higher levels of damping may beachieved, but which may diminish under constant temperature and load. Infuselage skin applications, for example, the temperature may berelatively constant although some variation may occur during cruiseconditions, and loads are also constant. Where the aircraft is exposedto a low level acoustic environment, the damping can be maximized undervariable conditions where temperature, static loads and acoustic loadsare fluctuating.

FIG. 3 effectively illustrates the temperature hysteresis of SEMmaterials relative to an imposed load. The temperature hysteresis can bemodified, for example, from −50 C to −5 C, up to +60 C to +120 C, byvarying the ratios of the metals forming the chosen alloy.

Ni—Ti is one desirable SEM that may be suitable for use as wire fibers46 in the hybrid composite structure 30. The temperature hysteresis andtransformation properties of Ni—Ti may be changed by adding appropriateamounts of Cu, Nb, Fe or Pt. Other suitable SEM materials which arecopper based, include Cu—Zn—Al, Cu—Al—Ni, and Cu—Al—Be. Further, somemonocrystalline alloys may be suitable for use since they exhibitsimilar, super-elastic properties. Other super-elastic, highly dampedalloys that may be used to fabricate the wire fibers 46 of the hybridfiber composite structure 30 include: Fe—C—Si, Al—Zn, Fe—Cr, Fe—Cr—Al,Co—Ni—Ti, Mg, Fe, Ni, Mg—Ni, Mn—Cu, Mn—Cu-AL, Cu—Zn—Al, Cu—Al, Ni,Ni—Ti, and Fe—Co.

The composite structure 30 illustrated in FIG. 2 comprises laminatedplies 32 of a fiber reinforced synthetic resin, such as a carbon fiberreinforced epoxy. The individual plies 32 may comprise a tape or fabriccontaining reinforcing fibers, such as carbon fibers. In the case oftape, the fibers unidirectional, whereas the fibers in the fabric may bebidirectional. The plies 32 may be oriented such that the reinforcingfibers are oriented at different angles relative to a referencedirection in order to increase stiffness and rigidity of the resultingcomposite structure 30. As is conventional in the art, the reinforcingfibers may be bundled in tows 42 (FIG. 3) which, in the case of afabric, may be knitted or woven to form a mat. In prepreg tape form, thetows 42 are normally arranged unidirectionally.

As shown in FIG. 4, the SEM wire fibers 46 may be interspersed with thereinforcing fibers 44 within the tows 42. As will be discussed later inmore detail, the diameter and number of SEM wire fibers 46 in each tow42 will depend on the particular application, and specifically thepercent volume of wire fibers 46 that are required to achieve a desiredlevel of damping for the application, with consideration given tostrength requirements. For example, the SEM wire fibers 46 may be,between 0.00005 inches and 0.005 inches in diameter. When tows 42 areemployed, multiple layers are of tape are placed on a lay-up tooling(not shown). The lay-up is then infused with a suitable resin such asepoxy, following which the plies are compacted and cured so that the SEMwire fibers 46 are held in the resin matrix, along with the reinforcingfibers 44.

As shown in FIG. 5, the SEM wire fibers 46 may be interspersed alongwith reinforcing fibers 44 in rows 50 that are held in a resin matrix 48as prepreg tape or fabric. Again, the size and number of the SEM wirefibers 46 will depend upon the level of damping that is desired andstrength requirements.

FIG. 6 shows another embodiment in which the SEM wires may be of smallerdiameter than the reinforcing fibers 44 and are placed within the spaces47 between adjacent ones of the fiber 44. Again, the SEM wire fibers 46are embedded in a surrounding resin matrix (not shown).

Reference is now made to FIGS. 7 and 8 which illustrate anotherembodiment of the composite structure 30 a having alternating groups ofplies 52, 54. The ply group 52 comprises fiber reinforced resin plies,whereas the ply groups 54 each comprise one or more layers of SEM fibers46 that are held in a resin matrix 48, as best seen in FIG. 7. Thus,rather than being integrated into plies containing reinforcement fibers44 as illustrated in FIGS. 4 and 5, the SEM wire fibers 46 are isolatedin separate ply groups 54.

The fuselage skin of modern composite aircraft is typically composed of8 to 32 plies of carbon/epoxy pre-preg clothe and tape. Each pre-pregtape layer is arranged in a different orientation, usually 0, 90, +/−45degrees. It is known through empirical data, that duringacoustic/vibration loading, the highest stresses are in layers are thenear the outer surfaces. A laminate for an aircraft skin may be composedof hybrid damped metallic fiber layers near the outer surfaces, andstandard carbon/epoxy plies near the interior layer. This layup wouldmaximize damping performance (and electrical and impact resistance) andminimize weight increase. The hybrid metallic pre-preg may typically be20-30% greater in areal weight than standard carbon/epoxy.Macro-mechanic analysis of the disclosed hybrid laminates indicates thatat least 10% to 20% of the plies may be hybrid plies in order to achievedamping levels of Loss Factor=0.03 to 0.05. This level of damping wouldrepresent a 2 to 5 times increase in the baseline damping of standardcarbon epoxy laminate.

Reference is now made to FIGS. 9 and 10 which illustrate SEM fibers 46embedded in an optional interlayer 56 contained within one or moresingle plies 52 a of fiber reinforced synthetic resin. The interlayer 56may be formed of a material that is relatively soft compared to thefiber reinforced resin in the plies 52, such as, without limitation, aviscoelastic material (VEM). VEM's encompass a variety of materialclassified as thermoplastics, thermoplastic elastomers or thermosets.The VEM in interlayer 56 may have a high loss tangent or ratio of lossmodulus to storage modulus, in order to provide the laminate structureformed by the plies 52 with the desired damping properties. The VEM inthe interlayer 56 may have a modulus that is approximately two or moreorders of magnitude less than the modulus of the resin used in the plies52.

In the illustrated example, the SEM wire fibers 46 are orthogonallyarranged as a mesh 58 that is embedded in the VEM interlayer 56. The SEMwire fibers 46 may have ends 46 a that extend beyond the interlayer 56and are anchored within the resin forming ply 52 a.

Reinforced prepreg plies of material containing a hybrid mixture ofreinforcing and SEM fibers may be produced using a fabrication lineillustrated in FIG. 11. SEM wire fibers 60 are fed from a roll 64 andare aligned and combined into a single layer or lamina, along withcarbon reinforcing fibers 62 supplied from a separate spool 66.Alignment of the fibers 60, 62 may be such that they form alternatingrows of carbon and SEM metal fibers that are then combined with aflexible film of resin 68 supplied from spool 70. The single layer ofhybrid fibers 60, 62 is pressed onto the resin film 68 by feed rollers72 and are passed over a heating element 74 which heats the resin film68 to its free-flowing temperature. Consolidation rollers 76 are used toimpregnate the melted resin film 60 into the hybrid fibers 60, 62 inorder to form a final, hybrid fiber prepreg tape 78 that is taken up ona spool 80.

FIG. 12 is a graph illustrating the effect of SEM metal fibers 46 onlaminate damping for an eight ply carbon fiber reinforced laminatestructure employing Ni—Ti wire fibers. Curve 82 demonstrates theincrease in the material loss factor with increasing percentage volumeof Ni—Ti wire fibers in the laminate. Curve 84 demonstrates thecorresponding decrease in the longitudinal modulus (Ell) of thelaminate. The graph of FIG. 11 demonstrates that good laminate dampingfor some aerospace applications such as commercial aircraft can beobtained by using 10 to 30 percent by volume of SEM fibers in thelaminate structure. The results shown in FIG. 11 may be predicted usingcomplex modulus macro-mechanic models to determine the effect of dampedSEM metal fibers on laminate damping.

Loss factor Lf is a property of a material which is a measure of theamount of damping in the material. The higher the loss factor, thehigher the damping. The loss factor of the material is sensitive to theload and temperature imposed on the structure. Typical reinforcedcomposite laminates have an inherent material loss factor (Lf) from0.001 to 0.01 depending upon temperature and stress levels. Whenincorporated into aircraft structures, laminate damping may be measuredat Lf=0.005 to 0.015. As shown in the graph of FIG. 11, the use of acomposite structure having hybrid reinforcing and SEM wire fibers mayincrease the inherent damping to Lf=0.03 to 0.10. Thus, a potentialincrease of 3 to 10 times over known composite laminates may be achievedthrough the use of embedded SEM wire fibers. In some applications, wherethe inherent laminate damping can be increased to Lf=0.03, there may notbe a need for the use of “add-on” damping devices.

Increasing the level of damping in composite structures for aerospacevehicles is practical where the increased damping significantly enhancesthe total system damping. The use of SEM wire fibers 46 to increasedamping may be evaluated in relation to the amount of damping initiallypresent in the structure, and the affect on the structure in terms ofthe weight, strength and stiffness of the resulting composite material.While the ultimate strength and stiffness of a composite structure maybe slightly reduced as a result of the use of SEM metallic fibers 46 toincrease damping, the SEM wire fibers 46 may cause the laminate tobecome tougher, in terms of fatigue and impact resistance. This may beparticularly important in aircraft where some structural parts aredesigned to achieve higher levels of stiffness or impact resistance, orfracture toughness.

It should be noted here that plies containing a hybrid mixture of fiberse.g. carbon and SEM wire fibers may be selectively used throughout thecomposite structure to achieve the desired level of damping, with a netreduction in weight compared to an all carbon fiber structure withadd-on damping. Thus, for example, in order to achieve 25-33% of SEMwire fibers by volume, every two plies out of eight plies or four pliesout of twelve plies may contain the SEM fibers in order to achieve adesired level of damping. Also, plies containing the hybrid fibermixture may be selectively used throughout a composite structure. Forexample, a composite laminate having hybrid fiber plies may be used onlyin the frames within a skin or in the caps of a hat, or in the shear-tieof a frame. In the case of a wing skin, the hybrid fiber plies may onlybe used in the inboard section of an upper wing skin, and only in themain spars.

Although the embodiments of this disclosure have been described withrespect to certain exemplary embodiments, it is to be understood thatthe specific embodiments are for purposes of illustration and notlimitation, as other variations will occur to those of skill in the art.

1-20. (canceled)
 21. A method of fabricating a damped compositestructure for aircraft, comprising the steps of: introducing shapememory alloy wire fibers into plies of fiber reinforced synthetic resin;and laminating the plies.
 22. The method of claim 21, wherein the stepof laminating the plies includes: laying up multiple fiber reinforcedplies of synthetic resin having the wire fibers introduced therein,compacting the plies of the layup, and curing the compacted layup. 23.The method of claim 21, wherein the step of introducing shape memoryalloy wire fibers includes forming tows of reinforcing fibers containingthe wire fibers.
 24. The method of claim 21, wherein the step ofintroducing shape memory alloy wire fibers includes forming a tapeincluding the wire fibers and reinforcing fibers.
 25. The method ofclaim 21, wherein the multiple fiber reinforced plies include a layer ofviscoelastic material and the wire fibers are embedded in theviscoelastic material
 26. The method of claim 25, wherein the layer ofviscoelastic material is contained in a layer of resin.
 27. The methodof claim 26, wherein the wire fibers extend from the layer ofviscoelastic material into the layer of resin.
 28. The method of claim22, wherein the step of layup up multiple fiber reinforced plies ofsynthetic resin comprises arranging the wire fibers as a mesh
 29. Themethod of claim 21, further comprising: carbon fibers reinforcing thematerial matrix, and
 30. The method of claim 29, wherein: the carbonfibers are arranged in a plurality of tows, and the wire fibers arecontained within the tows.
 31. The method of claim 21, wherein the shapememory alloy is selected from the group consisting of: Ni—Ti, Cu—Zn—Al,Cu—Al—Ni, Cu—Al—Be.
 32. The method of claim 31, wherein the shape memoryalloy includes one of: Nb, Fe and Pt.
 33. The method of claim 21,wherein the shape memory alloy is selected from the group consisting of:Fe—C—Si, Al—Zn, Fe—Cr, Fe—Cr—Al, Co—Ni—Ti, Mg, Fe, Ni, Mg—Ni, Mn—Cu,Mn—Cu-AL, CU-Zn—Al, CU-Al, Ni, Ni—Ti, and Fe—Co.
 34. The method of claim21, wherein the wire fibers exhibit an inherent material loss factorgreater than approximately 0.10.
 35. A method of fabricating a dampedcomposite structure for aircraft, comprising the steps of: providing amatrix material comprising a resin and a plurality of reinforcementfibers; and introducing a plurality of metallic shape memory alloy wirefibers into the matrix material.
 36. The method of fabricating thedamped composite structure of claim 35, wherein the plurality of matrixmaterial comprises a plurality of plies; and further comprisinglaminating the plurality of plies.
 37. The method of fabricating thedamped composite structure of claim 35, wherein the material matrixincludes a layer of viscoelastic material and the wire fibers areembedded in the viscoelastic material.
 38. The method of fabricating thedamped composite structure of claim 35, wherein the material matrixincludes a layer of resin and the layer of viscoelastic material iscontained in the layer of resin.
 39. The method of fabricating thedamped composite structure of claim 37, wherein the wire fibers extendfrom the layer of viscoelastic material into the layer of resin.
 40. Themethod of fabricating the damped composite structure of claim 35,further comprising arranging the wire fibers as a mesh.